Method for determining waveguide temperature for acoustic transceiver used in a gas turbine engine

ABSTRACT

A method for determining waveguide temperature for at least one waveguide of a transceiver utilized for generating a temperature map. The transceiver generates an acoustic signal that travels through a measurement space in a hot gas flow path defined by a wall such as in a combustor. The method includes calculating a total time of flight for the acoustic signal and subtracting a waveguide travel time from the total time of flight to obtain a measurement space travel time. A temperature map is calculated based on the measurement space travel time. An estimated wall temperature is obtained from the temperature map. An estimated waveguide temperature is then calculated based on the estimated wall temperature wherein the estimated waveguide temperature is determined without the use of a temperature sensing device.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a U.S. national stage application and claims thebenefit under 35 U.S.C. § 371 of International Application No.PCT/US2015/026784 filed on Apr. 21, 2015 and entitled METHOD FORDETERMINING WAVEGUIDE TEMPERATURE FOR ACOUSTIC TRANSCEIVER USED IN A GASTURBINE ENGINE which is incorporated by reference in its entirety and towhich this application claims the benefit of priority. InternationalApplication No. PCT/US2015/026784 is a continuation in part of copendingUnited States Patent Application entitled “PARAMETER DISTRIBUTIONMAPPING IN A GAS TURBINE ENGINE”, filed on Apr. 9, 2015, U.S.application Ser. No. 14/682,393, which is incorporated herein byreference in its entirety and to which this application claims thebenefit of priority. International Application No. PCT/US2015/026784also claims the benefit under 35 U.S.C. § 119(e) of U.S. ProvisionalPatent Application No. 61/983,044 entitled “TEMPERATURE DISTRIBUTIONMAPPING IN A GAS TURBINE COMBUSTOR”, filed on Apr. 23, 2014, which isincorporated herein by reference in its entirety and to which thisapplication claims the benefit of priority.

STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT

Temperature mapping portions of this invention were made with governmentsupport under contract DE-FC26-05NT42644 awarded by the U.S. Departmentof Energy. The government may have certain rights in the invention.

This application incorporates by reference the following co-pendingUnited States utility patent applications in their entirety as if fullyset forth herein:

“Nonintrusive Performance Measurement of a Gas Turbine Engine in RealTime”, filed on Jul. 28, 2014, Ser. No. 14/341,950; U.S. PatentPublication No. 2015/0260557;

“Nonintrusive Transceiver and Method for Characterizing Temperature andVelocity Fields in a Gas Turbine Combustor”, filed on Jul. 28, 2014,Ser. No. 14/341,924; U.S. Patent Publication No. 2015/0260611;

“Active Measurement Of Gas Flow Temperature, Including In Gas TurbineCombustors”, filed on Mar. 13, 2014, Ser. No. 14/207,741; U.S. PatentPublication No. 2015/0168230;

“Active Temperature Monitoring In Gas Turbine Combustors”, filed on Dec.18, 2013, Ser. No. 14/132,001; U.S. Patent Publication No. 2015/0168229;

“Multi-Functional Sensor System For Gas Turbine Combustion MonitoringAnd Control” filed on Dec. 18, 2013, Ser. No. 14/109,992; U.S. PatentPublication No. 2015/0168228;

“Temperature Measurement In A Gas Turbine Engine Combustor”, filed onMar. 14, 2013, Ser. No. 13/804,132; U.S. Patent Publication No.2014/0278200 and

“Gas Turbine Engine Control Using Acoustic Pyrometry”, filed on Dec. 14,2010, Ser. No. 12/967,148; U.S. Patent Publication No. 2012/0150413,U.S. Pat. No. 8,565,999.

This application also incorporates by reference in its entirety as iffully set forth herein U.S. Pat. No. 7,853,433, “Combustion AnomalyDetection Via Wavelet Analysis Of Dynamic Sensor Signals”, issued Dec.14, 2010.

BACKGROUND OF THE INVENTION

1. Field of the Invention

The invention relates to the mapping of a parameter in two-dimensionalspace and to active measurement of gas flow parameters, such as gas flowtemperature or velocity, in flow regions of gas turbine engines. Suchengines include, by way of example, industrial gas turbine (IGT)engines, other types of stationary gas turbine, marine, aero and othervehicular gas turbine engines. More particularly, embodiments disclosedherein disclose a method for determining a waveguide temperature for atleast one waveguide in order to include the effect of boundary wall andwaveguide temperatures on a temperature distribution of a temperaturemap, wherein the method includes calculating an estimated waveguidetemperature based on an estimated wall temperature and wherein theestimated wall temperature is determined without the use of atemperature sensing device.

2. Description of the Prior Art

Combustion turbines, such as gas turbine engines for any end useapplication, generally comprise a compressor section, a combustorsection, a turbine section and an exhaust section. In operation, thecompressor section inducts and compresses ambient air. The combustorsection generally may include a plurality of combustors for receivingthe compressed air and mixing it with fuel to form a fuel/air mixture.The fuel/air mixture is combusted by each of the combustors to form ahot working gas that may be routed to the turbine section where it isexpanded through alternating rows of stationary airfoils and rotatingairfoils and used to generate power that can drive a rotor. Theexpanding gas exiting the turbine section can be exhausted from theengine via the exhaust section.

Combustion anomalies, such as flame flashback, have been known to occurin combustion sections of gas turbine engines. Flame flashback is alocalized phenomenon that may be caused when a turbulent burningvelocity of the air and fuel mixture exceeds an axial flow velocity inthe combustor assembly, thus causing a flame to anchor onto one or morecomponents in/around the combustor assembly, such as a liner disposedaround the combustion chamber. The anchored flame may burn through thecomponents if a flashback condition remains for extended periods of timewithout correction thereof. Thus, flame flashback and/or othercombustion anomalies may cause undesirable damage and possibly evendestruction of combustion engine components, such that repair orreplacement of such components may become necessary.

The fuel/air mixture at the individual combustors is controlled duringoperation of the engine to maintain one or more operatingcharacteristics within a predetermined range, such as, for example, tomaintain a desired efficiency and/or power output, control pollutantlevels, prevent pressure oscillations and prevent flameouts. In a knowntype of control arrangement, a bulk turbine exhaust temperature may alsobe monitored as a parameter that may be used to monitor the operatingcondition of the engine. For example, a controller may monitor ameasured turbine exhaust temperature, and a measured change intemperature at the exhaust may result in the controller changing anoperating condition of the engine. In other known types of controlarrangements discrete pitot-static or multi hole pressure probes areutilized to determine gas flow velocity at specific locations, but gridarrays of such probes disrupt gas flow and introduce measurement errors.Due to such gas flow disruptions, grid arrays, when employed, havelimited numbers of widely spaced probes, which provide relatively coarsegas flow velocity distribution and profile information.

At present, there are several different types of sensors and sensingsystems that are being used in the industry for monitoring combustionand maintaining stability of the combustion process for engineprotection. For example, dynamic pressure sensors are being used forcombustion stability and resonance control. Passive visual (opticalvisible light and/or infrared spectrum) sensors, ion sensors and GeigerMueller detectors are used to detect flame on/off in the combustor,while thermocouples are being used for flashback detection. With respectto known combustion gas flow velocity (u) monitoring methods,pitot-static and multi hole pressure probes utilize differentialpressure techniques, hot wire probes utilize thermal anemometrytechniques, while Laser Doppler and Particle Image Velocimetry systemsutilize optical techniques to characterize gas flow velocities.Differential pressure and thermal anemometry instruments are intrusivepoint measurement devices that disturb local gas flow around theinstruments. Laser Doppler and Particle Image Velocimetry instrumentsrespectively provide non-intrusive point and 2- or 3-dimensionalnon-intrusive gas flow velocity measurement although they both requireparticle seeding of the flow. In addition, sophisticated laser basedmeasurements such as Filtered Rayleigh Scattering (FRS) and other suchlaser spectroscopy based techniques have been deployed to measure gasvelocity. However, these techniques are more complex than intrusivedifferential pressure or thermal anemometry instruments and require morespecialized training to implement in monitoring systems. Moreover, mostoptical techniques for velocity are geared towards laboratoryenvironments rather than in operative engines at power plant fieldsites. With respect to temperature (T) monitoring techniques, knownRaman Spectroscopy, Laser Induced Fluorescence (for both u and Tmonitoring), and Coherent Anti-Stokes Raman Spectroscopy (CARS) (forboth u and T monitoring) instrumentation systems are also intended forlaboratory environments, rather than for field use in fossil powergeneration equipment. Tunable Diode Laser Absorption Spectroscopy(TDLAS) instrumentation is used in some industrial power generationfield applications, such as for temperature measurement in boilers butthat instrumentation is extremely costly: approximately US $500,000 persystem. Other types of temperature measurement and combustion anomalydetection systems have had greater acceptance in power generationindustry field applications.

Particularly, U.S. Pat. No. 7,853,433 detects and classifies combustionanomalies by sampling and subsequent wavelet analysis of combustorthermoacoustic oscillations representative of combustion conditions withsensors, such as dynamic pressure sensors, accelerometers, hightemperature microphones, optical sensors and/or ionic sensors. UnitedStates Publication No. US2012/0150413 utilizes acoustic pyrometry in anIGT exhaust system to determine upstream bulk temperature within one ormore of the engine's combustors. Acoustic signals are transmitted fromacoustic transmitters and are received by a plurality of acousticreceivers. Each acoustic signal defines a distinct line-of-sound pathbetween a corresponding transmitter and receiver pair. Transmittedsignal time-of-flight is determined and processed to determine a pathtemperature. Multiple path temperatures can be combined and processed todetermine bulk temperature at the measurement site. The determined pathor bulk temperature or both can be utilized to correlate upstreamtemperature in the combustor. Co-pending U.S. utility patent applicationSer. No. 13/804,132 calculates bulk temperature within a combustor,using a so-called dominant mode approach, by identifying an acousticfrequency at a first location in the engine upstream from the turbine(such as in the combustor) and using the frequency for determining afirst bulk temperature value that is directly proportional to theacoustic frequency and a calculated constant value. A calibration secondtemperature of the working gas is determined in a second location in theengine, such as the engine exhaust. A back calculation is performed withthe calibration second temperature to determine a temperature value forthe working gas at the first location. The first temperature value iscompared to the back calculated temperature value to change thecalculated constant value to a recalculated constant value. Subsequentfirst temperature values at the combustor may be determined based on therecalculated constant value.

A need exists for techniques for creating real time, two-dimensionalmaps of temperature distribution in a flow region of a gas turbineengine based on estimates of average temperature along lines betweentransceivers and that include the effect of boundary wall and waveguidetemperatures on temperature distribution of the two-dimensionaltemperature map.

A further need exists in the art for an integrated gas turbine enginemonitoring and control system for measuring gas flow velocity,temperature and detecting a broad range of possible combustor failuresor, more satisfactorily, precursors to faults, during combustion,sharing common sensors and, if desired, a common controller.

Another need exists in the art for a gas turbine engine active velocityand temperature monitoring system that maps actual combustor velocityand temperature in real time without the need to obtain referencetemperatures from other locations within the engine, such as known bulktemperature systems that back calculate combustor temperature based ontemperature measurements obtained in the engine exhaust system.

An additional need exists for an active gas flow velocity andtemperature monitoring system that shares sensors commonly used withcombustion turbine monitoring and control systems, so that activevelocity and temperature monitoring can be integrated within themonitoring and control system.

A further need exists for a technique for providing real-timetemperature information in a plane transverse to a gas flow in a turbineengine for controlling the engine.

Another need exists for a technique for controlling a gas turbinecombustor based on average temperature measurements along lines in aplane transverse to the combustor flow.

SUMMARY OF THE INVENTION

A method is disclosed for determining a waveguide temperature for atleast one waveguide used in conjunction with a transceiver utilized forgenerating a gas turbine temperature map. The transceiver generates anacoustic signal that travels through a measurement space in a hot gasflow path defined by a wall such as in a combustor. The method includescalculating a total time of flight wherein the total time of flightincludes a travel time through the measurement space and a travel timethrough the waveguide. The waveguide travel time is then subtracted fromthe total time of flight to obtain a measurement space travel time. Themethod also includes calculating a temperature map based on themeasurement space travel time and then obtaining an estimated walltemperature from the temperature map. Further, the method includescalculating an estimated waveguide temperature based on the estimatedwall temperature wherein the estimated waveguide temperature isdetermined without the use of a temperature sensing device.

In addition, a method for determining a waveguide temperature for atleast one waveguide used in conjunction with a transceiver utilized forgenerating a gas turbine temperature map is disclosed. The transceivergenerates an acoustic signal that travels through a measurement space ina hot gas flow path defined by a wall such as in a combustor. The methodincludes dividing the wall into a plurality of boundary portions whereina transceiver and waveguide is associated with each boundary portion.The method also includes calculating a total time of flight for theacoustic signal wherein the total time of flight includes a travel timethrough the measurement space and a travel time through the waveguide.The waveguide travel time is then subtracted from the total time offlight to obtain a measurement space travel time. In addition, themethod includes calculating a temperature map based on the measurementspace travel time and then obtaining from the temperature map anestimated temperature for each boundary portion. Further, the methodincludes calculating an estimated waveguide temperature for eachboundary portion wherein the estimated waveguide temperature isdetermined without the use of a temperature sensing device.

The respective objects and features of the present invention may beapplied jointly or severally in any combination or sub-combination bythose skilled in the art.

BRIEF DESCRIPTION OF THE DRAWINGS

The teachings of the present invention can be readily understood byconsidering the following detailed description in conjunction with theaccompanying drawings, in which:

FIG. 1 is a perspective cross-sectional view of a gas turbine engineillustrating implementation of a system for determining combustor gasflow active velocity and temperature measurement, in accordance withembodiments of the invention;

FIG. 2 is a cross-sectional view of a gas turbine combustorincorporating an embodiment of a monitoring system for determiningcombustor gas flow active velocity and temperature measurement, inaccordance with embodiments of the invention;

FIG. 3 is a cross-sectional view of the system of FIG. 2, taken along3-3 thereof, in accordance with aspects of the invention;

FIG. 4 is a block diagram of an embodiment of a controller forimplementing embodiments of the present invention in the monitoringsystem for determining combustor gas flow active velocity andtemperature measurement, in accordance with embodiments of theinvention;

FIG. 5 is a schematic perspective view of exemplary sonic sensor arraysused by the gas flow monitoring system to measure gas flow velocity in agas turbine combustor, in accordance with embodiments of the invention;

FIG. 6 is an exemplary schematic representation of gas flow velocity inthe turbine combustor of FIG. 5 in the line-of-sight between acousticsensors 32B and 34C;

FIG. 7. is a cross-sectional slice A of the gas flow velocity of FIG. 6taken along 7-7 thereof, which corresponds to the line-of-sight betweenacoustic sensors 32B and 34C;

FIG. 8 is a composite gas flow velocity profile of the respectivevelocities measured by the gas flow velocity monitoring system, inaccordance with embodiments of the invention;

FIG. 9 is a schematic perspective view of exemplary sonic sensor arraysused to measure gas flow temperature in a gas turbine combustor, inaccordance with embodiments of the invention;

FIG. 10 is a flow chart illustrating implementation of an embodiment ofthe methods for measuring gas flow velocity and temperature activemeasurement in a gas turbine combustor, in accordance with embodimentsof the invention; and

FIG. 11 is a flow chart illustrating implementation of an embodiment ofthe method for measuring active gas flow velocity, in accordance withembodiments of the invention.

FIG. 12 is a schematic illustration of a gas turbine engine showingsensor installations in several alternative regions, in accordance withembodiments of the invention.

FIG. 13 is a schematic illustration of a system for mapping flowparameters in a gas turbine engine region, in accordance withembodiments of the invention.

FIG. 14A is a schematic depiction of a bilinear representation of aparameter along a path, in accordance with embodiments of the invention.FIG. 14B is a schematic representation of a two-dimensional space withtwo bilinear representations of path profiles, in accordance withembodiments of the invention.

FIG. 15 is a flow chart showing a technique for mapping a parameterbased on average path values, according to embodiments of the invention.

FIG. 16 is a diagram showing parameter profiles for a single path,according to embodiments of the invention.

FIG. 17 is a flow chart showing a technique for mapping a parameterbased on average path values, according to embodiments of the invention.

FIG. 18 is a schematic view of a two dimensional space showingmeasurement paths and a grid segment, according to embodiments of theinvention.

FIG. 19 is a flow chart showing a technique for mapping a parameterbased on average path values, according to embodiments of the invention.

FIG. 20 is an exemplary cross sectional view of a plurality oftransceivers each having a waveguide wherein the transceivers eachgenerate an acoustic signal that defines a plurality of acoustic pathstravelling through a measurement space of a hot gas flow path.

FIG. 21 depicts a boundary portion of a wall that defines the hot gasflow path wherein the wall is divided into a plurality of virtualboundary portions each having an associated transceiver and waveguide.

FIG. 22 depicts a boundary portion of the wall that is not divided andthat includes transceivers each having a waveguide.

FIG. 23 is a flowchart for illustrating a method for determining awaveguide temperature in accordance with embodiments of the invention.

FIG. 24 is a chart depicting an exemplary waveguide temperature plot foreach waveguide shown in FIG. 20.

FIGS. 25-30 depict a sequence of exemplary temperature maps thatillustrate temperature convergence for a waveguide as the number ofiterations increase.

To facilitate understanding, identical reference numerals have beenused, where possible, to designate identical elements that are common tothe figures.

DETAILED DESCRIPTION

After considering the following description, those skilled in the artwill clearly realize that the teachings of the invention can be readilyutilized for active acoustic velocity and pyrometry-based gas flowvelocity and temperature measurement. Embodiments of the invention areused for monitoring of gas turbine combustors, including industrial gasturbine (IGT) combustors by incorporating them into the combustionmonitoring and control system by addition of an acoustic transmitter oracoustic transceiver that transmits sound waves through gas flow in aline-of-sight with a plurality of acoustic sensors, such as dynamicpressure sensors. For velocity measurement, sound transmissiontime-of-flight that is directed generally transversely through the gasflow path is measured by the controller and correlated with gas flowvelocity along the line-of-sight. The gas flow velocity determinationincludes compensation for impact of the thermodynamically interrelatedtemperature, gas constant and speed of sound influences on the firsttime-of-flight, in order to determine absolute gas flow velocity.

In an integrated acoustic pressure-based sensor and monitoring/controlsystem embodiment, the controller correlates velocity and, if desired,absolute active path temperatures simultaneously with acoustictransmission and time-of-flight analysis techniques. Where velocity andtemperature are measured simultaneously the absolute active pathtemperature is utilized to compensate for the aforementionedthermodynamic influences on gas flow absolute velocity. Alternatively inother embodiments the speed of sound influence on the firsttime-of-flight is utilized to determine absolute gas flow velocityrather than absolute active path temperature. In such embodiments,compensation for the speed of sound in the velocity monitoring isaccomplished by substituting for the first transmitters a set of firsttransceiver/transducers that are capable of transmitting and receivingacoustic signals, and generating output signals and substituting for thefirst sensors a set of second transducers that are capable oftransmitting and receiving acoustic signals and generating outputsignals. Acoustic signals are transmitted and received from the first tothe second transducers and time-of-flight is determined. A reverseacoustic signal is transmitted from the second to the first transducersand the reverse time-of-flight is determined. The respective first andfirst reversed acoustic signals times-of-flight are used to determinethe speed of sound c. The determined speed of sound c is then utilizedfor determination of the actual gas flow velocity.

In embodiments of the invention active velocity or activevelocity/temperature measurements are used as monitoring parameters forgas flow in a combustion monitoring and control system that can identifyand classify gas flow anomalies (e.g., combustion anomalies), forexample by using wavelet or Fourier analysis techniques. Someembodiments of the methods and system incorporate one or more acousticdynamic pressure transceiver/transducer combination transmitter/sensorsthat are selectively oriented or arrayed in sequential axial planarpositions within the combustor. Known transceiver/transducer componentdesigns and their related controller components have been used reliablyand cost effectively in the past in power generation field service. Byreconfiguring those types of known components into the gas flow controland monitoring systems of the present invention combustion turbine andother combustion power generation equipment can be monitored andcontrolled with simpler instrumentation hardware configurations thatprovide detailed active gas flow velocity and temperature distributioninformation useful for precise combustion control.

Monitoring and Control System Structure

Referring to FIGS. 1 and 2 an exemplary industrial gas turbine engine 10is shown. The exemplary engine 10 includes a compressor section 12, acombustor section 14, a turbine section 16, and an exhaust section orsystem 18. The combustor section 14 includes a plurality of combustors20. Each combustor 20 has a combustion shell 22 and a cover plate 24.The combustor liner or basket 26 and transition duct 27 define a passagefor conveying hot working gas that flows in the direction F to theturbine section 16. The system of the present invention is operable withknown combustor geometry gas turbine engine designs, including can,can-annular or annular construction combustors in stationary land-basedor vehicular applications.

During operation of the engine 10, compressed air from the compressorsection 12 is provided to the combustor section 14 where it is combinedwith fuel supplied by fuel injection system 28 in the combustors 14. Thefuel/air mixture is ignited to form combustion products comprising thehot working gas. It may be understood that combustion of the fuel andair may occur at various axial locations along the passage through thecombustor liner or basket 26 and the transition duct 27 to the inlet ofthe turbine section 16. The hot working gas is expanded through theturbine section 16 and is exhausted through the exhaust section/system18.

Referring to FIGS. 1 and 2, in accordance with an aspect of theinvention, a combustion monitoring and control system 29 is provided,which can identify and classify combustion anomalies and activelycontrol the gas turbine combustion process within one or more of theengine 10 combustors 20. In this regard, the engine 10 may include maycomprise one or more of the monitoring and control system(s) 29: e.g.,one system 29 for each combustor 20, or a single system 29 may serviceeach combustor 14 of the engine 10. Similarly, clusters of combustors 20may be served by one system 29, with other cluster(s) being served byother systems. Thus the consolidated monitoring system for an engine 10can determine deviations between respective combustors and compare theirrelative performance no matter what engine combustor structure ororientation is employed by the engine design: whether a stationary,land-based turbine engine or a vehicular engine for aero, marine or landvehicular applications.

As shown in FIGS. 2, 3, 5 and 9 the system 29 includes an array of aplurality of known acoustic transceiver/transducers 32A-H and 34A-H thatare capable of transmitting and receiving acoustic oscillation wavesalong exemplary the line-of-sight paths shown in dashed lines in FIGS. 5and 9. The transceiver/transducer arrays 32, 34 are capable ofgenerating respective sensor output signals indicative of combustionthermoacoustic oscillations in each respective monitored and controlledcombustor 20. Other system embodiments can be constructed with at leasttwo, but preferably more acoustic sensors, whether functionally part ofa transceiver component or as a stand-alone component. Acousticfrequencies and amplitudes sensed by those acoustic sensor portions ofthe transceivers are generated as a result of combustion events in theworking combustion gas, defining acoustic sources that occur within thecombustor 20 hot gas path. The monitoring and control system 29 isconfigured to transform the sensed thermoacoustic oscillationinformation into a form that enables the occurrence of combustionanomalies of interest to be discerned. As such, flame flashback eventsand other types of combustion anomalies of interest may be detected andextracted from sensed thermoacoustic oscillations in the combustor 14that are monitored by the transceiver/transducer/sensors positioned inand/or around the combustor 14. Depending upon the system 29configurations and application, the acoustic sensors comprise anycombination of one or more of a dynamic pressure sensor, a microphone,an optical sensor or an ionic turbine inlet sensor. Pressure sensorssense the amplitudes of thermoacoustic oscillations in the combustor 20as well as pulsation frequencies. A high temperature microphone may beutilized to measure acoustic fluctuations in the combustor 14. Anoptical sensor may be utilized to measure a dynamic optical signalwithin the combustor 20. An ionic sensor may be utilized to measuredynamic ionic activity within the combustor 20.

An exemplary acoustic sensor array shown schematically in FIGS. 2, 3, 5and 9 comprises transceiver/transducers 32A-H and 34A-H that function asat least one acoustic transmitter that transmits in turn to at least oneand preferably a plurality of the dynamic pressure sensors in the array.The transceiver/transducers 32, 34 are arrayed axially and radiallywithin the combustor 20 by known mounting structures and methods, suchas J tubes or rakes, within the combustor shell 22 proximal thecombustor basket or liner 26, and/or proximal the transition 27 junctionwith the turbine section 16. In FIG. 3 the sensors areradially/circumferentially arrayed transceivers 34A-34H that are capableof transmitting and receiving acoustic oscillation waves along theline-of-sight paths similar to the transceivers 32A-H shown in dashedlines in FIG. 9. Other types of known sensors, such as individualthermocouple temperature sensors or thermocouple arrays may be employedwithin the gas turbine engine. For example in FIG. 3 thermocouple 36measures combustion temperature in the combustor 20. While exemplarythree-dimensional annular combustion flow paths and axially spaced,two-dimensional circular-annular transceiver/transducer arrays are shownin the figures, other combustion flow path and array orientations may beutilized, in practicing embodiments of the invention, including square-or rectangular-shaped geometries.

As shown in greater detail in FIGS. 3 and 4, the monitoring and controlsystem 29 comprises a known controller 40, coupled to thetransceiver/transducers 32, 34, that is capable of correlating sensoroutput signals with gas flow velocity and combustion temperature in amonitoring section 42 and conducting combustion dynamics analysis of thecombustion process in an analysis section 44. The monitoring section 42and dynamic analysis 44 section outputs are utilized by the gas turbinecontrol system 46 that can send control signals to other gas turbinecontrols subsystems, including industrial gas turbine (IGT) controlssubsystems, such as the fuel injection system 28, in order to unload orshut down the engine 10 in response to changes in monitored combustionconditions within the combustor 20.

Referring to the exemplary controller 40 embodiment shown in FIG. 4, itincludes one or more processors 50, system memory 52 and input/outputcontrol devices 54 for interfacing with the associated engine 10controls, such as the fuel injection control system 28, and the acoustictransceiver/transducer 32, 34 acoustic transmitters and sensors 32 (orfunctionally equivalent performing separate discrete transmitters andreceiver sensors), networks, other computing devices, human machineinterfaces for operator/users, etc. The controller 40 may also includeone or more analog to digital converters 56A and/or other componentsnecessary to allow the controller 40 to interface with the transceivers32, 34 and/or other system components to receive analog sensorinformation. Alternatively, and/or additionally, the system 29 mayinclude one or more analog to digital converters 56B that interfacebetween the transceivers 32, 34 (or functionally equivalent performingseparate discrete transmitters and receiver sensors) and the controller40. As yet a further example, certain transceivers 32, 34 may have ananalog to digital converter 56C integral therewith, or are otherwiseable to communicate digital representations of sensed informationdirectly to the controller 40.

The processor(s) 50 may include one or more processing devices such as ageneral purpose computer, microcomputer or microcontroller. Theprocessors 50 may also comprise one or more processing devices such as acentral processing unit, dedicated digital signal processor (DSP),programmable and/or reprogrammable technology and/or specializedcomponent, such as application specific integrated circuit (ASIC),programmable gate array (e.g., PGA, FPGA).

The memory 52 may include areas for storing computer program codeexecutable by the processor(s) 50, and areas for storing data utilizedfor processing, e.g., memory areas for computing wavelet transforms,Fourier transforms or other executed mathematical operations used tooperate the monitoring and control system 29, as described more fullyherein below. As such, various aspects of the present invention may beimplemented as a computer program product having code configured toperform the detection of combustion engine anomalies of interest,combustion dynamics and engine control functions as set out in greaterdetail herein.

In this regard, the processor(s) 50 and/or memory 52 are programmed withsufficient code, variables, configuration files, etc., to enable thecontroller 40 to perform its designated monitoring and controlfunctions. For example, the controller 40 may be operatively configuredto sense thermoacoustic conditions, analyze thermoacoustic conditionsbased upon inputs from one or more transceiver/transducers 32, 34,control features of the engine 10 in response to its analysis, and/orreport results of its analysis to operators, users, other computerprocesses, etc. as set out in greater detail herein. Thus, all of thedynamic output signals originating from transceiver/transducers 32, 34may be communicated to a single processor 50. In this implementation,the single processor 50 will process the sensor dynamic output signalsusing the data analysis and control functions described in greaterdetail herein, such that it appears as if the results are computed in agenerally parallel fashion. Alternatively, more processors 50 can beused and each processor may be utilized to process one or moretransceiver/transducers 32, 34 dynamic signals, e.g., depending forexample, upon the computation power of each processor.

Monitoring and Control System Operation

The concepts of acoustic temperature and velocity measurements are bothbased on creating a sonic wave, listening to it across the gas streamand finding an average speed of sound across a given path, which is thendescriptive for the gas velocity or velocity/temperature. FIGS. 10 and11 are flow charts illustrating graphically exemplary operation of amonitoring and control system 29 embodiment of the invention thatactively monitors and measures both gas flow velocity and temperatureusing acoustic measurement methodologies. The thick solid and dottedline operational blocks relate to previously described combustiondynamics analysis 42 (solid block), temperature monitoring anddetermination 44 and gas turbine control 46 functions (including by wayof example IGT control functions) that are performed within thecontroller 40. In step 100 sensor signals generated by the sensorcomponents within the transceiver/transducers 32A-H, 34 A-H are read. Instep 110 amplitudes of one or more of the sensor signals are compared topreviously established alarm limits. For example in IGT applications thestep 120 low frequency dynamics (LFD) below 100 Hz are of importancebecause of potential resonance influence at the 50 Hz or 60 Hz enginerotational speed. Other frequency bands of interest are intermediatefrequency dynamics (IFD) between approximately 100-500 Hz and highfrequency dynamics (HFD) above 500 Hz. If an alarm limit is exceeded thecontroller 40 sends a control command, for example to the fuel injectionsystem 28, to unload or shut down the engine 10 in step 400.

If an alarm limit is not exceeded in step 110, then frequency analysisfor dynamics is performed in anomaly detection portion of the combustiondynamics analysis sub system. An exemplary description of how to performanomaly detection is in U.S. Pat. No. 7,853,433 that is incorporatedherein by reference. The sampled high speed dynamic pressure signal isobtained from the sensors in step 130 and time divided into segments instep 140. In step 150 the time-frequency divided sample segments areanalyzed using the wavelet analysis technique described in U.S. Pat. No.7,853,433. Alternatively, a known Fourier spectral analysis thatconverts the time segments into frequency space, analyzes dominantfrequencies by identifying the peak frequencies and their respectiveamplitudes, and identifies amplitudes exceeding defined thresholds. Ifit is determined that a combustion anomaly or anomalies have occurred instep 160 the combustor temperature as determined in the temperaturemonitoring and determination subsystem 44 is compared with the anomalyinformation obtained by the Fourier or wavelet analysis techniques, orboth. In step 180 the anomaly classification as a flame on, flame out orflashback is made in conjunction with the passive or path temperatureinformation obtained from the temperature monitoring and determinationsubsystem 44. For example in a gas turbine flameout the combustortemperature drops off dramatically. Conversely in a flashback scenariothe combustor temperature rises dramatically upstream within thecombustor 14. When the anomaly determination is made in step 180appropriate control signals to unload or shut down the engine are madein the engine control system 46.

The temperature monitoring and determination subsystem 44 may comprisepassive temperature determination utilizing the passive acoustic methoddescribed in United States Patent Application “Temperature Measurementin a Gas Turbine Engine Combustor, filed on Mar. 14, 2013, Ser. No.13/804,132, incorporated by reference herein, and/or real time actualpath temperature determination within the combustor 14. Real time actualpath temperature is determined by adaptation of the 2-D planar acousticpyrometry technique for gas turbine exhaust system temperaturedetermination described in United States Patent Publication No.US2012/0150413 (also incorporated by reference herein) or by a 3-Dtechnique that determines one or more path temperatures between thesensor arrays 32/34 of FIG. 5, that is further described in greaterdetail herein.

In the passive temperature determination method, sampled high speeddynamic pressure signals from the transceiver/transducers 32/34, such asobtained in step 130 are analyzed for dominant modes in step 200.Combustor temperature is calculated based on frequency using the passiveacoustic method in step 210. The passive value is calibrated with areference temperature value in step 220 in order to obtain an activetemperature value within the combustor 14. The calibrated passivetemperature value determined in step 220 is utilized in step 230 todetermine the bulk mean temperature of the combustion gas in step 230.The reference temperature value used in step 220 may be obtained fromone or more thermocouples 36 in the combustor or thermocouples locatedin the exhaust system 18 (not shown). The reference temperature valuemay be an actual path temperature measured in the exhaust system 18, asdescribed in United States Patent Publication No. US2012/0150413 or areal time path temperature measured in the combustor 14 that isdetermined in steps 300-330.

The 2-D real time path temperature is measured by transmitting one ormore acoustic signals in an acoustic transceiver/transducer 32, 34 orother discrete transmitter, such as in the 2-D planar pattern shown forthe (n=8+transceiver/transducers 32A-H in FIG. 9. For example,transceiver/transducer 32A transmits a signal that is received by theremaining (n−1) transceiver/transducers 32B-H and the time-of-flight foreach line-of-sight path is determined. However, at least one, preferablytwo or more sensor elements in the remaining transceiver/transducers32B-H receive the acoustic signal(s) in step 310. Preferably in practiceseveral transceiver/transducers (transmit and receive acoustic signals)circling one plane such that the paths between all transceivers form agrid with desired coarseness which results in the spatial resolution ofthe temperature measurement. For example, for a cylindrical combustorthe transceivers could be equally spaced around the periphery as shownin FIGS. 3 and 9. These could be either fired sequentially (one at atime) or simultaneously with disjoint sound patterns that can be readilydifferentiated. For sequential firing one transceiver is creating soundswhile all remaining transceivers record it to estimate the travel timefor the respective paths. Each of these line-of-sight paths representsan average temperature along that path. The average temperatures overdifferent paths are combined to a two-dimensional map shown in FIG. 9,using a known computer tomography technique.

The 2-D time-of-flight sound data are converted to gas temperature usingactive acoustics in step 320, such as by utilization of the methodsdescribed in the aforementioned United States Patent Publication No.US2012/0150413 that is incorporated by reference herein. The real timepath temperature that is determined in step 330 is the localized activetemperature value along the line-of-sight transmission path. A pluralityof active temperature values measured along different acoustic paths byperforming the steps 300-330 can be utilized to determine the combustor14 bulk temperatures, alone or in parallel with the dominant frequencypassive acoustic method of steps 200-230. While a single path activetemperature measurement between a single transmitter 30 and acousticsensor 32 provides useful control information, arraying a plurality oftransceiver/transducers 32, 34 selectively in any axial, circumferentialacid/or radial pattern or combinations thereof within a combustor 14(see, e.g., FIG. 2, 3, 5, or 9) or in a series of combustors 14facilitates active real time two- or three-dimensional combustiontemperature monitoring within the gas turbine engine 10.

The 2-D or 3-D real time path temperature determined in steps 300-330can be utilized as an input for other monitoring and control functions,with or without one or more of the combustion dynamics analysis 42,passive temperature monitoring and determination 44 and control 46functions described in the exemplary integrated monitoring and controlsystem 29 described herein. For example combustor turbine inlettemperature (FIT) can be actively monitored in real time and used as acontrol parameter for the combustion process. The combustion active pathtemperature determined in steps 300-330 can be utilized to control thefuel/air mixture in the combustor 14 via the fuel injection system 28.The real time path active temperature can be utilized as an input foractive actual gas flow velocity measurement in an industrial gas turbinecombustor or in other types of gas flow environments.

Embodiments of the present invention measure 3-D gas flow velocityand/or gas flow temperature by correlation with sonic time-of-flightalong a line-of-sight sonic pathway between axially spaced, transverselyoriented sonic transmitter and sensor (or transceiver/transducersincorporating the sensors and transmitters), so that the line-of-sightalong the pathway is oriented transverse, as opposed to parallel to thegas flow path. In order to determine gas flow absolute velocity, thetime-of-flight data are corrected or compensated for thermodynamicinfluences on gas temperature, gas constant and speed of sound. As notedabove gas temperature along a line of sight can be determined using thereal time active path temperature or temperature independently obtainedfrom another measurement device (e.g., thermocouple 36). Alternativelylocalized speed of sound c can be determined by measuring bi-directionaltime-of-flight (i.e., forward/downstream transmission andreverse/upstream transmission). The aforementioned thermodynamicinfluences are governed by the known equation:c(x,y,z)=(γ·R·T)^(1/2)Where:c(x,y,z) is the isentropic speed of sound;γ is specific heat ratio;R is the gas constant; andT is the gas temperature.Therefore, once the speed of sound along a path is known, the averagepath temperature and absolute velocity can be determined utilizingembodiments of the invention further described herein.

For accurate absolute velocity or temperature measurement, two planes oftransceiver/transducers 32, 34 are oriented in axially spaced, opposedrelationship within the gas flow, as shown in FIG. 5. The twotransceiver/transducer planes 32, 34 are preferably apart byapproximately the same order of magnitude as the diameter (circular) orwidth (square or rectangular) of the monitored gas flow geometry. Thatis, the axial distance between the two planes should be determinedaccording to the geometry and scale of the interrogated environment aswell as the anticipated or possible ranges of gas flow gas constant,temperature and velocity.

For gas flow velocity estimation, the gas flow is measured axially andtransverse to the flow direction. For example, whentransceiver/transducer 32A in plane Z_(I) fires or transmits a signal,all transceiver/transducers 34B-H in plane Z_(II) that are notparallel-aligned with the signal firing sensor will be listening,thereby creating several paths across the gas flow (n−1 paths for nsensors). The signal transmitting/receiving firing process continuessequentially with the second transceiver/transducer 32B on plane Z_(I)firing to the remaining (n−1) transceiver/transducers 34A and 34C-H,which receive that transmitted signal. The transmitted signal firingwill continue on with the consecutive transceivers firing and creatingn−1 paths for each firing. In the embodiment of FIG. 5, having 8transceivers/transducers in each of the two axially spaced arrays thereare a total of 64 paths in three dimensions. In addition, to alleviatethe directional ambiguity of the velocity (to identify reverse flows andperhaps turbulence fluctuations in reverse direction) the same processwill be repeated with transducer/transceivers 34 in plane Z_(II) firingand transceiver/transducers in plane Z_(I) receiving the reversedirection transmitted acoustic signal, assuming that the gas flowtemperature is already known. Instead of transmitting/firing acousticsignals sequentially from each transceiver/transducer, a sound patternwith a slightly different acoustic signature can be transmitted fromeach respective transceiver/transducer 32A-H, 34A-H simultaneously,which shortens measurement time Referring to steps 500 and 510 of thegas flow velocity measurement method flow chart of FIG. 11, once alltransceiver/transducers in planes Z_(I) and Z_(II) have fired and thetransmitted acoustic signals have been received by the opposing plane oftransversely aligned transceivers/transducers, the process preferablyrepeats continually in real time while a 3-D velocity map u isconstructed from the spatially distributed line-of-sight acoustic paths,using known 3-D tomographic mapping techniques, such as those utilizedin medical or industrial computed tomography systems. The velocityinformation is extracted and mapped, as shown in FIG. 8. Similarly, a3-D temperature map T can be constructed utilizing the time of flightdata, as will be described in greater detail herein.

After all of the transceiver/transducers 32, 34 in a planar array havefired acoustic signals the respective line-of-sight flow pathtime-of-flight data are used to derive absolute velocity in the gas flowpath in step 560, once corrected for the thermodynamic effects oftemperature, gas constant and the speed of sound, as described ingreater detail below. Flow velocity measurement accuracy potentiallydecreases as flow velocity approaches the speed of sound, assumingconstant gas temperature in the velocity measurements. Flow velocitybelow a Mach number of approximately 0.5 is not believed to impactvelocity measurement significantly. Therefore it is preferable, but notrequired, that measured flow velocities should be smaller than half ofthe local speed of sound that is measured. This method can accuratelymeasure high temperature gas flows, including turbine engine gas flows,despite relatively high absolute velocities, because the local speed ofsound increases with temperature.

Once acoustic time-of-flight data are available, they are used by themonitoring and control system 29 or other remote monitoring system todetermine velocity along their respective acoustic paths in accordancewith the remainder of the steps of FIG. 11. Referring to FIGS. 6 and 7,information sound propagation is linearly affected by the gas flow.Relative gas flow velocity for a given temperature, gas constant andspeed of sound is determined by the known equation:

$t_{BC} = {\int_{B}^{C}{\frac{1}{{c\left( {x,y,z} \right)} + {{\overset{->}{p}}_{BC} \cdot {\overset{->}{u}\left( {x,y,z} \right)}}}{ds}}}$

-   -   Where:    -   t_(BC) is the time-of-flight from the first transmitter B to the        first sensor C;    -   c is the speed of sound in the gas flow for the temperature and        gas constant;    -   {right arrow over (p)}_(BC) is the unit vector along the first        line of sound path A between B and C; and    -   {right arrow over (u)}(x,y,z) is velocity vector in the gas        flow.

The exemplary planar slice along the line-of-sound path A shows asimplified flow pattern. Referring again to the flow chart of FIG. 11,the relative gas flow velocity is corrected for thermodynamictemperature, gas flow and speed of sound influences, in order to deriveabsolute velocity in step 560. If the path temperature is available(step 520) its influence on the speed of sound can be corrected by knowntomography methods, in order to derive the gas flow absolute velocityalong the line-of-sound path. If the path temperature is not available,times-of-flight for forward (steps 500, 510) and reverse (steps 530,540) acoustic signal transmission are acquired and used to extract thespeed of sound without effect of the gas velocity in accordance with thefollowing equations. The reverse time-of-flight fromtransducer/transceiver C to transducer/transceiver B is determined bythe following equation, similar to that for the forward or downstreamdirection set forth above:

$t_{CB} = {\int_{C}^{B}{\frac{1}{{c\left( {x,y,z} \right)} + {{\overset{->}{p}}_{BC} \cdot {\overset{->}{u}\left( {x,y,z} \right)}}}{ds}}}$

The forward and reverse times-of-flight are added in accordance with thefollowing equation:

${t_{BC} + t_{CB}} = {\int_{B}^{C}{\frac{2 \cdot {c\left( {x,y,z} \right)}}{{c\left( {x,y,z} \right)}^{2} + {{\overset{->}{p}}_{BC} \cdot {\overset{->}{u}\left( {x,y,z} \right)}^{2}}}{ds}}}$

Given that the square of the speed of sound c is much greater than thesquare of the gas flow velocity u, the equation is reduced to:

${t_{BC} + t_{CB}} \approx {\int_{B}^{C}{\frac{2}{c\left( {x,y,z} \right)}{ds}}}$

-   -   where:    -   t_(BC) is the time of flight from the first        transceiver/transducer B to the second transceiver/transducer C;    -   t_(CB) is the time of flight from the second        transceiver/transducer C to the first transceiver/transducer B;    -   c is the speed of sound in the gas flow for the temperature and        gas constant;    -   {right arrow over (p)}_(BC) is the unit vector along the first        line of sound path; and    -   {right arrow over (u)}(x,y,z) is the velocity vector in the gas        flow.

The speed of sound c determined in step 550 of FIG. 11 is then used tocorrect the downstream time-of-flight data for that speed of sound instep 560. The corrected downstream time-of-flight data are used todetermine gas flow absolute velocity in step 570. Where the pathtemperature T along a line-of-flight is not known, the same speed ofsound c determined in step 550 is utilized in some embodiments of theinvention to determine T, using the previously described isentropicspeed of sound relationship c(x,y,z)=(γ·R·T)^(1/2), as γ, R and c(x,y,z)are now known. In a similar manner to the path velocity determinationspreviously described, once all the path temperatures T are known fromeach receiver/transmitter unit back and forth, there will be 64(assuming exemplary 8-sensor case) iso-temperature lines in3-dimensions. Then using known 3-D tomographic mapping techniques, the3-dimensional temperature distribution is mapped.

Advantageously the active acoustic temperature and velocity measurementsare performed simultaneously in real time, thus mapping both gas flowtemperature (3-D or alternatively the 2-D mapping of FIG. 9) and 3-D gasflow velocity (FIG. 8). An exemplary acoustic signal transmission andreception timing sequence to perform simultaneous velocity andtemperature measurement is to emit an acoustic signal with atransceiver/transducer on a first array plane (e.g., 32A at Z_(I)). Thecorresponding transversely oriented transceivers/transducers on anaxially spaced opposed second plane 34B-H at Z_(II)) receive the signalfor velocity processing and/or temperature processing, if 3-Dtemperature measurement is utilized. If only 2-D temperature measurementis utilized the remainders of the transceiver/transducers on the firstarray plane (e.g., 32B-H at ZI) receive the signal for temperatureprocessing. As previously noted the transmission and receiving processalso can be accelerated by utilizing unique signal transmission patternsfor each transceiver/transducer. There are tradeoffs associated with useof 2-D or 3-D temperature measurement. Where 3-D temperature measurementtechniques are utilized, accuracy of both temperature and velocity mapmay not be the most desired in case of gas velocities of Mach 0.3 orabove as the approximation shown in the equation

${t_{BC} + t_{CB}} \approx {\int_{B}^{C}{\frac{2}{c\left( {x,y,z} \right)}{ds}}}$may be less accurate in those velocities ranges, because there are noindependently determined temperature reference values. However,independent temperature T reference values may be determined, using apair of axially separated 2-D acoustic signal sets and two individualacoustic temperature maps determined with the respective 2-Dtime-of-flight signal sets. The 2-D temperature maps are in turninterpolated to create a volumetric temperature map. This volumetric mapwill be used to provide the temperature values T utilized in theisentropic speed of sound equation, along with the known gas constant Rand specific heat ratio to extract speeds of sound c. The speed of soundis then used to extract the velocity vectors u(x,y,z). Once the velocityvectors are extracted the velocity components can be mapped, eliminatingthe limitation of below Mach 0.3 gas velocities inherent in thepreviously descried 3-D velocity and temperature mapping methods.

Combustor active gas flow velocity or velocity/temperature monitoringutilizing the system and method embodiments described herein with arraysof commonly utilized acoustic sensors is believed to provide fastervelocity and temperature change response than known velocity andtemperature monitoring systems. In accordance with embodiments of theinvention one array of commonly utilized, reliable acoustictransceiver/transducer sensor-transmitters or arrays of separatediscrete acoustic sensors and transmitter pairs can be placed in acombustion flow path under field conditions and monitored to provideactive, real time simultaneous velocity and temperature data and anomalydetection that are all useful for monitoring and control of combustionpower generation equipment, such as industrial gas turbines.

Mapping Parameter Distributions

A parameter map in a two-dimensional or three-dimensional space hasnumerous uses in the design, diagnosis and control of machinery. Forexample, a temperature or velocity map of a region of the gas path isuseful in diagnosing and accurately measuring the performance of a gasturbine engine. The map may, for example, be a temperature map in thevicinity of the combustor flame or may be a turbine inlet temperaturemap in the region exiting the combustor. Simple temperature maps arepresently created using thermocouple temperature rakes and temperatureprobes mounted on the first row vane to obtain measurements from which acrude linear temperature profile may be fit. Those short-term, intrusivemethodologies provide crude profiles based on sensor locations but donot provide a spatially resolved map of temperatures in real time thatone could utilize to control a gas turbine or to understand temperaturedistribution during a new engine design validation process.

Presently described are techniques for using acoustic or other signalsthat are transmitted and received in a region of a gas turbine engine.Many flow regions of a gas turbine engine may be of interest in the useof the presently described techniques, and several exemplary regions aredepicted in the schematic diagram of the gas turbine engine 1200 shownin FIG. 12. An inlet temperature map 1211 of a gas turbine inlet 1210may be created using the described techniques and using acoustic sensors1212 arranged circumferentially around a planar region of the inlet. Acombustor temperature map 1221 may be created to show temperaturedistributions in regions of the combustors 1220. Depending on the areaof interest, the sensors 1222 may be arranged around a plane through aprimary combustor flame zone or a turbine inlet (combustor exit). Athree-dimensional velocity map 1231 of gas flow through a turbinediffuser 1230 may be constructed using information from sensors 1232arranged circumferentially around multiple planar regions in thediffuser. An exhaust temperature map 1241 may be created using sensors1242 to show temperature distribution in a two-dimensional region of theturbine exhaust 1240. One skilled in the art will recognize that thedescribed techniques may be implemented with other sensor arrangementsand other gas turbine engine regions to yield additional usefulparameter maps.

An advanced methodology of tomographic principles is utilized to extracta spatially resolved map in real time using a few dozen individual pathsof the signals. As described above, the sensors may be acoustic sensorsand speed of sound information on each path is processed to estimate theaverage temperature over that path length. At each time interval,representations of each of the paths containing average temperatureinformation are tomographically mapped into the spatial distribution ofthe temperatures at the measurement time and then updated for subsequentmeasurement times. Information from the resulting temperature map may beused in the engine control algorithm or to maintain safety of engineoperation and low emission levels.

In embodiments as discussed above and as shown in FIG. 13, thetransmitters and sensors 1310 are circumferentially distributed around across section of the hot gas path of one or more turbine regions 1305.The sensors and receivers may, in some embodiments, be acoustictransceivers (sender/receiver combinations) arranged in a plane throughthe combustor where those transceivers will send and capture acousticsignals in real-time. While the disclosure discusses the sensors andreceivers with reference to acoustic sensing techniques, one skilled inthe art will understand that the sensors and receivers may alternativelyutilize laser-based tunable diode laser absorption spectroscopy oranother measurement technology to determine an average temperature alonga line path in the combustor. In the case of acoustic transceivers, thetransmitted signals are used to determine an average speed of sound,which is used to estimate an average temperature.

In the case of laser based tunable diode laser absorption spectroscopy,several techniques for temperature measurement are possible. Pathaveraged temperatures may be measured by probing two differentabsorption lines for the same species while sweeping the laser acrossthe absorption spectrum. Laser absorption by gases across the plane incertain infrared wavelength bands is proportionate to speciesconcentration and temperature and can be solved to provide averaged pathtemperatures along each line. Alternatively, the Full Width at HalfMaximum (FWHM) of the probed absorption line may be related to theDoppler line width of the species. Other laser-based or othertemperature measurement techniques may be used without departing fromthe scope of the disclosure.

In the case of temperature mapping, a tomographic mapping module 1315converts a plurality of average path temperatures into a temperature map1320 at each time interval that the temperatures are sampled. Thetwo-dimensional or three dimensional map includes high spatialresolution isotherms and provides more valuable information thanseparate average path temperature estimates for interpretation of theengine health as well as for inputs to the engine control algorithm. Thetemperature map 1320, together with derived combustion qualityinformation 1325, is transmitted to an engine control unit 1330 thatutilizes that information to control the combustors and/or the gasturbine engine.

While described with reference to the construction of a temperature mapfrom average path temperature estimates, the techniques described hereinmay be used to construct other two dimensional or three dimensional mapsfrom average values of paths. For example, the estimated averagevelocities along transmitter receiver paths may be used to construct atwo-dimensional or three dimensional map of local velocities using thesimilar methods.

Described herein are several different techniques for mapping andextracting spatial information about a parameter from a set of pathaveraged lines in a region. Those techniques include a polynomialapproximation method, a basis function method and a grid optimizationmethod. Each of those techniques is described in turn below. Whileseveral of the descriptions refer to an exemplary embodiment in which agas flow temperature is measured, one skilled in the art will recognizethat the described techniques are applicable to mapping other parameterswhere average values along linear paths are available.

Polynomial Approximation Technique:

One way of accomplishing the task of converting a plurality of averagepath temperatures to a temperature map is to approximate the temperatureprofiles along each path with a polynomial and then, through aniterative process as shown in the flow chart 1500 of FIG. 15, to adjustthe parameters of each polynomial to minimize errors. To do this, eachpath, such as path 1405 shown in FIG. 14A, is initially assigned, inoperation 1510 (FIG. 15), a temperature function that includes scalefactors and reflects the estimated average temperature along the path.In the bi-linear profile 1400, temperatures increase linearly from theend points 1415, 1420 (i.e., the transmit and receive points at thechamber walls) to form a peak 1425 yielding a profile similar to thecross section of a tent. The initial maximum temperature is at the peakand the minimum (wall) temperatures are at either end. The distance 1430from the transmit end point 1415 to the peak 1425 is determined by amidpoint scale parameter. The midpoint scale parameter may be initiallyset to 50% of the path length by default.

The peak height 1410 defines the initial maximum temperature along thepath 1405 at the peak 1425. Initially, a peak height scale factor may beset to the value of two times the average path temperature.

The end points 1415, 1420 may be assumed to be constant and are held ata constant level by the algorithm relative to a wall temperaturevariable. The wall temperature variable may be selected in several ways.In one example, a fixed value is manually entered. In another example, apercentage of the average path temperature, or the minimum pathtemperature, is used. In other embodiments, an actual sensor such as ahigh temperature thermocouple is used to input a wall temperature signaldirectly into the algorithm.

Bilinear integration is performed from the transmit point 1415 (walltemperature) to the mid-point 1425 (measured path temperatureiteratively computed scale factor) and back down to the receive point1420 (wall temperature).

The estimated path profile 1400 is then plotted, in operation 1520, intoa two dimensional grid representing the planar area of the chamber 1401,as shown in FIG. 14B. That process is repeated for each path, such asexemplary second path 1450. When all temperatures have been plotted ontothe grid, the grid contains a sparse representation of the temperaturemap. As a result, there are missing points on the grid in the open areasbetween paths. In those open areas, a curve smoothing technique such asa Bezier function is used to transform the set of known points into apolynomial approximation of the actual temperature curve at discretepoints on the grid. After the grid is smoothed, line integrations areperformed along line paths, in operation 1530, and compared to measureddata in operation 1540. The result of that comparison is used to adjustthe scale factors in block 1570 for the next iteration to minimize theerror between measured and estimated temperatures. The iterative processmay be repeated according to decision operation 1550 between 3 and 20times to produce a surface with the least error when compared to theoriginal measured average path temperatures.

The iterations are ended in operation 1560 when the errors are below apreset maximum. While there is no specification for absolute error oraccuracy of the generated isothermal map, an average path error valuecan be calculated to provide a confidence factor for any given map. Theaverage path error will range from 0 to 3-4% for a typical system innormal operation. If the mapping plane and the physical factors havebeen optimally selected, this can be used for a very accurate spatialmap from the path averaged temperature information.

Basis Function Technique:

Another technique for converting a plurality of average pathtemperatures to a two dimensional temperature map is the use of basisfunctions. In general, every continuous function in a function space maybe represented by a linear combination of basis functions. In thepresently described technique, portrayed in the flow chart 1700 of FIG.17, the temperature map is represented by a linear combination of2-dimensional basis functions derived from thermocouple measurements.

The 2-dimensional basis functions are extracted, as shown in operation1710, from a large database of thermocouple temperature measurements orother parameter measurements, using a statistical procedure such asprincipal component analysis (PCA). The boundary conditions arefixed/constant for the 2-dimensional basis functions based on a manuallyentered fixed value, or on measured wall temperatures. The techniquefinds weights for the basis functions that minimize error on measuredtimes of flight. In embodiments, the weights are found iteratively.

In one embodiment, there are K two-dimensional basis functions derivedfrom the thermocouple measurements. There are furthermore I acousticpaths i and t_(i) mean temperatures, one estimated from each time offlight measurement along an acoustic path. For each basis function andpath, the path temperatures are sampled (operation 1720) into a vectoralong a length D of the path. For each path 1, the sampled temperaturesare collected in a D×K matrix X_(i), as illustrated by the matrix 1600of FIG. 16, which shows, for a single path, six temperature profiles forsix basis functions K₁-K₆, each sampled at five locations D₁-D₅. Forexample, for the basis function K₁ shown in matrix 1600, fivetemperature samples are illustrated for five sampled points D1 throughD5 along the path. Similar matrices are created for each of the otherpaths. The goal is to find (operation 1730) the combination of basisfunctions that best represents the mean temperatures measured by thetimes of flight. The weighted combination of basis functions is given bythe weight vector a which can be found as:

$\arg\;{\min\limits_{a}{\sum\limits_{i = 1}^{I}{{{{\frac{1}{D}a^{T}X_{i}\underset{\_}{1}} - t_{i}}}^{2}.}}}$

Grid Optimization Technique:

In an embodiment utilizing acoustic signals, the estimated average pathtemperatures may be converted to a two dimensional temperature map usinga grid optimization technique shown in the flow chart 1900 of FIG. 19,in which the 2-dimensional map is segmented into multiple grid segments(operation 1910), such as the grid segment 1820 shown in FIG. 18. Thegoal is to estimate the value (temperature or velocity) in each gridsegment. A grid segment is defined as a region bounded by the horizontaland vertical lines of the grid. The speed of the acoustic signalthroughout each grid segment is assumed to be uniform.

The distance covered by each acoustic path in traversing each gridsegment is initially calculated, at operation 1920. In the example shownin FIG. 18, the distances within the grid segment that are traversed bypaths 1831, 1833, 1835 and 1837 are determined. Since the total time offlight for each path and the distance traversed by the path through eachgrid segment is known, the time taken to travel through each gridsegment along a path can be calculated by solving following system ofequations:

$\begin{bmatrix}t_{1} \\t_{2} \\\ldots \\t_{n}\end{bmatrix} = {\begin{bmatrix}d_{11} & d_{12} & \ldots & d_{1m} \\d_{21} & d_{22} & \ldots & d_{2m} \\\ldots & \ldots & \ldots & \ldots \\d_{n\; 1} & d_{n\; 2} & \ldots & d_{nm}\end{bmatrix}\begin{bmatrix}x_{1} \\x_{2} \\\ldots \\x_{m}\end{bmatrix}}$where n is an index of paths, in is an index of grid segments, t is thetime of flight for a given path, x is a coefficient of each grid segmentcorresponding to a reciprocal of the speed of sound in that grid segment

$\left( \frac{1}{velocity} \right)$and d is the distance traversed by a path n through each grid segment m.

A boundary condition is applied on the grid segment coefficientscorresponding to the boundaryx _(w) =cwhere w is a grid segment index corresponding to the wall and c is aconstant derived from the wall temperature.

To reduce the search space and limit the solved result to withinacceptable ranges, upper and lower bounds for the coefficients areimposed based on the physical reality. For example, the temperature ofeach grid segment may be restricted to a value greater than roomtemperature and less than 1000° C.

The number of grids segments to be solved may greatly exceed the numberof path equations, making a large number of solutions possible. In thatcase, additional optimization criteria such as minimizing the differencebetween speeds of sound in neighboring grid segments are imposed toobtain a smooth map which is also close to reality.

After computing a value for a speed of sound for each grid segment(operation 1930), temperature values are estimated for the gridsegments, and a temperature map may be constructed (operation 1940)using those values.

In embodiments, the above techniques may be employed in controlling agas turbine engine using temperature data. Once the two-dimensionaltemperature map is computed, it is available in real-time for thecomputation of information useful in controlling the engine. Forexample, referring to FIG. 13, the tomographic mapping module 1315 maycompute the bulk mean temperature (average temperature of the map), thedistribution of temperature in the plane (as a goodness function or as aprofile), and a difference of temperature between different baskets ifthe combustion system is a can or can annular combustion system. Thatinformation is then provided to the engine control unit 1330, which isintelligently programmed to control engine parameters for optimum engineperformance (safety, performance and emissions).

As previously described, transceiver/transducers 32A-H and 34A-H (seeFIG. 5) transmit and receive acoustic signals that define acousticpaths. Time of flight measurements along corresponding acoustic pathsare then used to generate a temperature map such as temperature map 1320shown in FIG. 13. Each acoustic signal is transmitted and received bytransceiver/transducers 32A-H and 34E-H through a waveguide associatedwith a transceiver/transducer.

Referring to FIG. 20, an exemplary cross sectional view is shown of aplurality of acoustic paths 2000 travelling through a measurement space2010 of a hot gas flow path 2020. For purposes of illustration, the hotgas flow path 2020 is located in a combustor 20 although it isunderstood that the hot gas flow path 2020 may be located in the turbinediffuser 1230 (see FIG. 12) or an exhaust stack. FIG. 20 depicts aplurality of transceiver/transducers (i.e. transceivers) each having awaveguide, for example transceivers 2040A-2040F having correspondingwaveguides 2030A-2030F, respectively. The transceivers 2040A-2040F onlymeasure temperature and gas flow in a volume between the transceivers2040A-2040F, i.e., the measurement space 2010. However, a temperature ata solid boundary that defines the hot gas flow path 2020 is notmeasured. Further, a temperature inside the waveguides 2030A-2030F isalso not measured. In an embodiment, the solid boundary may be a wall2050 of the combustor 20.

In order to generate a temperature map, an assumed or selectedtemperature is used for the wall 2050 and waveguides 2030A-2030F. It hasbeen found by the inventors herein that the wall 2050 and waveguide2030A-2030F temperatures are important parameters that affect thetemperature distribution and temperature values indicated on atemperature map. However, the selected or assumed temperatures that areused are frequently erroneous, resulting in inaccurate temperature mapsfor a hot gas flow that has not been previously characterized.

A temperature map includes temperature information near a wall thatforms a hot gas flow path. In accordance with embodiments of theinvention, waveguide temperatures are calculated based on thetemperature near the wall 2050 that is indicated by an initialtemperature map. In an embodiment, a linear relationship is establishedbetween a wall temperature value and waveguide temperature although itis understood that nonlinear relationships may be used. Referring toFIG. 21, a boundary portion 2060 of the wall 2050 is divided into aplurality of virtual boundary portions 2070A-2070H (eight exemplaryboundary portions are shown) thereby thrilling a boundary portion index.Each boundary portion 2070A-2070H is associated with correspondingtransceivers 2040A-2040H that include the corresponding waveguides2030A-2030H, respectively, such that the waveguides 2030A-2030H arelocated adjacent or relatively close to a corresponding boundary portion2070A-2070H. The temperature of each waveguide 2030A-2030H is governedby the temperature of its associated boundary portion 2070A-2070H inaccordance with Eq. (1):tempAtWaveGuide(tr_i)=0.75×estimated_wall_value(matching_boundary_portion)+30K  (1)wherein “tr_i” corresponds to a transducer/receiver 2040A-2040H numberassociated with a boundary portion 2070A-2070H, “tempAtWaveGuide(tr_i)”is a temperature for a waveguide 2030A-2030H corresponding to thetransducer number “tr_i” and “K” is the Kelvin temperature scalealthough it is understood that other temperatures scales may be used.The “estimated_wall_value(matching_boundary_portion)” is an estimatedtemperature for the boundary portion 2070A-2070H associated with thetransducer. This enables calculation of a temperature for each boundaryportion 2070A-2070H and thus a corresponding waveguide temperature basedon each boundary portion 2070A-2070H, resulting in a plurality ofwaveguide temperatures as shown in FIG. 24. As will be described, Eq.(1) may be used to calculate an initial waveguide temperature.

In another embodiment, the wall 2050 is not divided as shown in FIG. 22and a uniform value for temperature is assumed for each waveguide2030A-2030H. A temperature is calculated of a complete boundary portion2080. The temperature of all the waveguides 2030A-2030H is given by Eq.(2):tempAtWaveGuide=0.75×mean_estimated_wall_value+30K  (2)wherein “tempAtWaveGuide” is the temperature of all the waveguides2030A-2030H, “mean_estimated_wall_value” is the estimated meantemperature for the complete boundary portion 2080 and “K” is the Kelvintemperature scale although it is understood that other temperaturesscales may be used.

Referring back to FIG. 20, a total time of flight for an acoustic signalis the sum of the amount of time for the acoustic signal to travelthrough a waveguide 2030A-2030H and the amount of time for the acousticsignal to travel through the measurement space 2010. The time taken totravel through a waveguide 2030A-2030H is subtracted from the total timeof flight and the remaining time is then used to generate a temperaturemap. Waveguide temperature is an important parameter that affects thetime of flight of the acoustic signal. Calculation of an updatedwaveguide temperature via Eq. (1), for example, causes a correspondingchange in the time taken to travel in the measurement space 2010. Thisin turn causes a remapping of the temperature map. Once a newtemperature map is generated, a new wall temperature is obtained fromthe new temperature map which is then used to calculate a new orestimated waveguide temperature again using Eq. (1). It is understoodthat Eq. (2) may be used instead of Eq. (1) to update the waveguidetemperature.

The waveguide temperature is continually updated, and new temperaturemaps are generated, if a difference between an estimated waveguidetemperature and an initial waveguide temperature, calculated immediatelybefore the estimated waveguide temperature, is greater than apredetermined temperature difference i.e., a temperature differentialthreshold. If the difference between the estimated waveguide temperatureand the initial waveguide temperature is less than or equal to thetemperature differential threshold, the temperature of a boundaryportion and the temperature of its associated waveguide have convergedand the process stops.

Referring to FIG. 23, a flowchart 2090 is shown which illustrates amethod for determining a waveguide temperature in accordance withembodiments of the invention. At the beginning of the method, an assumedinitial waveguide temperature 2105 for each waveguide is initially used.In an embodiment, the initial waveguide temperature 2105 is set to roomtemperature, although it is understood that other temperatures may beused. At Step 2100, the total time of flight is the sum of the amount oftime for the acoustic signal to travel through a waveguide 2030A-2030Hand the amount of time for the acoustic signal to travel through themeasurement space 2010. The waveguide travel time is subtracted from thetotal time of flight to provide a travel time through the measurementspace 2010 at Step 2110. At Step 2120, a temperature map is calculatedbased on the travel time through the measurement space 2010. Next, awall temperature is extracted from the temperature map at Step 2130. AtStep 2140, an estimated waveguide temperature of each waveguide2030A-2030H is calculated based on Eq. (3):estWaveGuide(tr_i)=0.75×estimated_wall_value(matching_boundary_portion)+30K  (3)

wherein “tr_i” corresponds to a transducer 2040A-2040H number associatedwith a boundary portion 2070A-2070H, “estWaveGuide(tr_i)” is anestimated temperature for a waveguide 2030A-2030H corresponding to thetransducer number “tr_i” and “K” is the Kelvin temperature scalealthough it is understood that other temperatures scales may be used.The “estimated_wall_value(matching_boundary_portion)” is an estimatedtemperature for the boundary portion 2070A-2070H associated with thetransducer.

At Step 2150, if the difference between the estimated waveguidetemperature of each waveguide 2030A-2030H, (i.e. “estWaveGuide(tr_i)”)from Eq. (3) and an assumed initial waveguide temperature, or an initialwaveguide temperature equal to a previously calculated estimatedwaveguide temperature, is greater than a temperature differentialthreshold, the initial waveguide temperature is set equal to theestimated waveguide temperature of each waveguide 2030A-2030H(“estWaveGuide(tr_i)”) at Step 2155. Thus, the initial waveguidetemperature and the estimated waveguide temperature of each waveguideoccur at different times during the method. The method then returns toStep 2110 and Steps 2120, 2130 and 2140 are repeated to perform anotheriteration of the method and obtain a new estimated waveguide temperatureof each waveguide 2030A-2030H (“estWaveGuide(tr_i)”).

If the conditions in Step 2150 are not met, i.e. the difference betweenthe estimated waveguide temperature, i.e. Eq. (3) and an initialwaveguide temperature is less than or equal to the temperaturedifferential threshold, the temperature of a boundary portion2070A-2070H of the wall 2050 and the temperature of its associatedwaveguide 2030A-2030H have converged and the process stops. This alsoindicates that the wall temperature and waveguide temperatures arerelatively constant. In an embodiment, the temperature differentialthreshold is approximately 5° C. Aspects of the method may beimplemented as an algorithm or computer program for use in thepreviously described tomographic mapping module 1315, computer system orother computing device.

Referring to FIG. 24, a chart 2160 is shown depicting an exemplarywaveguide temperature plot 2170 for each waveguide 2030A-2030F. Eachtemperature plot 2170 is obtained during iteration of the method. Fromthe chart 2160, it can be seen that each waveguide temperature begins toconverge (i.e. the difference between the estimated waveguidetemperature and an initial waveguide temperature as previously describedbecomes increasingly smaller) as the number of iterations of the methodincrease thereby improving the accuracy of the indicated waveguidetemperature. In this example, the waveguide temperatures converge inregion 2180 of the chart 2160 after approximately 30 iterations therebyindicating accurate temperatures. In an embodiment, a fast mappingalgorithm may be used that iterates multiple times per measurement toallow convergence of the waveguide temperatures without substantiallyaffecting real time capabilities of the overall system 29. This enablesfaster convergence of the waveguide temperatures in the event there is alarge difference between the initial and estimated waveguidetemperatures and faster generation of an approximate temperature map.Once the differential waveguide temperature reduces to a selectedthreshold the temperature map may be solved more accurately.

FIGS. 25-30 depict a sequence of exemplary temperature maps 2190, 2200,2210, 2220, 2220 and 2240, respectively, that illustrate temperatureconvergence for a waveguide as the number of iterations increase.Temperature maps 2190-2240 include temperature regions 2250, 2260, 2270,2280 whose arrangement and indicated temperature changes as the numberof iterations increase, thus changing temperature distribution of theassociated temperature map. An initial temperature map 2190 is shown inFIG. 25 wherein waveguide temperatures of 35, 38, 36, 30, 30, 31° C. areinitially estimated for waveguides 2030A-2030F, respectively. FIGS.26-29 illustrate iterations of the method wherein the waveguidetemperatures are updated and the temperature maps change accordingly ineach figure. FIG. 30 depicts a final temperature map 2240 wherein thedifference between an estimated waveguide temperature and an initialwaveguide temperature is less than or equal to a temperaturedifferential threshold, i.e. the temperatures converge as previouslydescribed. In FIGS. 25-30, additional iterations (up to 45) wereperformed in order to show stable algorithm convergence.

Estimation of boundary and waveguide conditions is repeated only ifthere are significant changes so as to limit computational cost andimpact on the overall system 29. That is, once the boundary andwaveguide condition converge, the algorithm keeps track of the changingboundary conditions to maintain accuracy of the total map. In practice,a complete temperature map is calculated once the boundary condition andwaveguide temperature converge. In an embodiment, a learned history of arelationship between waveguide temperature and a temperature map may beused to enhance the estimation of waveguide temperatures. For example, arecord or data of the relationship between waveguide temperature and atemperature map from a similar gas turbine or experimental engine havingsimilar temperature characteristics may be used to enhance estimation ofwaveguide temperatures. In addition, an estimated waveguide temperatureobtained from a previous temperature mapping solution may be used inorder to increase the speed of convergence of the waveguidetemperatures. Further, constraints on possible mapping solutions such aslimiting temperature ranges such as a temperature range for thewaveguide or other temperature ranges, allowing only a linearcombination of basis maps and others may be used to increase accuracyand speed of temperature mapping.

The current invention provides a method for automatically determiningboundary and waveguide conditions and extends mapping to use theboundary and waveguide conditions. In particular, waveguide and boundarytemperatures are iteratively calculated in real time without need fortheir actual measurement (i.e. without the need to measure temperatureby using a temperature sensing device or probe) in order to obtain atrue temperature map. The current invention reduces the need forinstrumenting waveguides and walls and other areas to obtain boundaryconditions. Further, the current invention increases the accuracy oftemperature, velocity and mass flow measurement of hot gas flows.

Although various embodiments that incorporate the teachings of thepresent invention have been shown and described in detail herein, thoseskilled in the art can readily devise many other varied embodiments thatstill incorporate these teachings. The invention is not limited in itsapplication to the exemplary embodiment details of construction and thearrangement of components set forth in the description or illustrated inthe drawings. However, the various aspects of the present inventiondescribed more fully herein may be applied to other instances where aprofile map of values in a region is determined based on average valuesalong linear paths through the region. The invention is capable of otherembodiments and of being practiced or of being carried out in variousways. While acoustic and laser sensors are discussed, other measurementtechniques may be used. Also, it is to be understood that thephraseology and terminology used herein is for the purpose ofdescription and should not be regarded as limiting. The use of“including,” “comprising,” or “having” and variations thereof herein ismeant to encompass the items listed thereafter and equivalents thereofas well as additional items. Unless specified or limited otherwise, theterms “mounted,” “connected,” “supported,” and “coupled” and variationsthereof are used broadly and encompass direct and indirect mountings,connections, supports, and couplings. Further, “connected” and “coupled”are not restricted to physical or mechanical connections or couplings.

What is claimed is:
 1. A method for determining a waveguide temperaturefor at least one waveguide used in conjunction with a transceiver thatgenerates an acoustic signal that travels through a measurement space ina hot gas flow path defined by a wall, comprising: calculating, by aprocessor, a total time of flight for the acoustic signal wherein thetotal time of flight includes a travel time through the measurementspace and a travel time through the waveguide; subtracting, by theprocessor, the waveguide travel time from the total time of flight toobtain a measurement space travel time; calculating, by the processor, atemperature map based on the measurement space travel time; obtaining,by the processor, an estimated wall temperature from the temperaturemap; and calculating, by the processor, an estimated waveguidetemperature based on the estimated wall temperature wherein theestimated waveguide temperature is determined without the use of atemperature sensing device; controlling, by a controller, a gas turbinecombustion process based on the estimated waveguide temperature.
 2. Themethod according to claim 1, wherein the wall is divided into aplurality of boundary portions each having an associated transceiver andadjacent waveguide wherein a temperature of each boundary portion isused to calculate a corresponding estimated waveguide temperature. 3.The method according to claim 2, wherein the estimated waveguidetemperature is linearly related to the estimated wall temperature. 4.The method according to claim 1, wherein the method is performediteratively thereby providing iterative estimation of waveguidetemperatures based on estimated temperature maps.
 5. The methodaccording to claim 4, wherein a fast mapping algorithm is used thatiterates multiple times per measurement to enable waveguide temperatureconvergence without substantially affecting real time capabilities of anassociated system.
 6. The method according to claim 1, wherein thetemperature differential threshold is approximately 5° C.
 7. A methodfor determining a waveguide temperature in a gas turbine, comprising thesteps of: (a) providing at least one transceiver that generates anacoustic signal that travels through a measurement space in a hot gasflow path defined by a wall in the gas turbine, wherein the at least onetransceiver includes a waveguide; (b) calculating, by a processor, atotal time of flight for the acoustic signal wherein the total time offlight includes a travel time through the measurement space and a traveltime through the waveguide; (c) subtracting, by the processor, thewaveguide travel time from the total time of flight to obtain ameasurement space travel time; (d) calculating, by the processor, atemperature map based on the measurement space travel time; (e)obtaining, by the processor, an estimated wall temperature from thetemperature map; (f) calculating, by the processor, an estimatedwaveguide temperature based on the estimated wall temperature whereinwhen a difference between the estimated waveguide temperature and aninitial waveguide temperature is greater than a temperature differentialthreshold, steps (c)-(e) are repeated wherein a converged waveguidetemperature is determined when a difference between the estimatedwaveguide temperature and an initial waveguide temperature is less thanor equal to the temperature differential threshold and wherein theestimated waveguide temperature is determined without the use of atemperature sensing device; (g) utilizing a learned history of arelationship between waveguide temperature and a temperature map toenhance estimation of waveguide temperatures; controlling, by acontroller, a combustion process of the gas turbine utilizing theestimated waveguide temperature; unloading or shutting down the gasturbine when an anomaly is determined by the processor from theestimated waveguide temperature.
 8. The method according to claim 7,wherein the wall is divided into a plurality of boundary portions eachhaving an associated transceiver and adjacent waveguide wherein atemperature of each boundary portion is used to calculate acorresponding estimated waveguide temperature.
 9. The method accordingto claim 8, wherein the estimated waveguide temperature is linearlyrelated to the estimated wall temperature.
 10. The method according toclaim 7, wherein the method is performed iteratively thereby providingiterative estimation of waveguide temperatures based on estimatedtemperature maps.
 11. The method according to claim 10, wherein the afast mapping algorithm is used that iterates multiple times permeasurement to enable waveguide temperature convergence withoutsubstantially affecting real time capabilities of an associated system.12. The method according to claim 7, wherein the temperaturedifferential threshold is approximately 5° C.
 13. A method fordetermining a waveguide temperature for at least one waveguide used inconjunction with a transceiver that generates an acoustic signal thattravels through a measurement space in a hot gas flow path defined by awall, comprising: (a) dividing, by a processor, the wall into aplurality of boundary portions wherein a transceiver and waveguide isassociated with each boundary portion; (b) calculating, by theprocessor, a total time of flight for the acoustic signal wherein thetotal time of flight includes a travel time through the measurementspace and a travel time through the waveguide; (c) subtracting, by theprocessor, the waveguide travel time from the total time of flight toobtain a measurement space travel time; (d) calculating, by theprocessor, a temperature map based on the measurement space travel time;(e) obtaining, by the processor, from the temperature map an estimatedtemperature for each boundary portion; and (f) calculating, by theprocessor, an estimated waveguide temperature for each boundary portionwherein the estimated waveguide temperature is determined without theuse of a temperature sensing device; controlling, by a controller, a gasturbine combustion process based on the estimated waveguide temperature.14. The method according to claim 13, wherein the estimated waveguidetemperature is linearly related to the estimated wall temperature. 15.The method according to claim 13, wherein the method is performediteratively thereby providing iterative estimation of waveguidetemperatures based on estimated temperature maps.
 16. The methodaccording to claim 15, wherein a fast mapping algorithm is used thatiterates multiple times per measurement to enable waveguide temperatureconvergence without substantially affecting real time capabilities of anassociated system.
 17. The method according to claim 13, wherein thetemperature differential threshold is approximately 5° C.
 18. The methodaccording to claim 13, wherein the estimated waveguide temperature isobtained by a convergence of waveguide temperatures and an estimatedwaveguide temperature obtained from a previous temperature mappingsolution is used to increase a speed of convergence of the waveguidetemperatures.
 19. The method according to claim 18, wherein convergenceoccurs after approximately 30 iterations.